1. Field of the Invention
The present invention relates to the rotor system of a helicopter excluding the upper controls, and to a method of fabricating component parts of the rotor system from composite material.
The rotor system presented is a fully articulated rotor system, although the invention has application in other rotor system configurations.
2. Prior Art
The traditional articulated rotor system configuration for operational helicopters with three or more rotor blades has been:
(1) a machined metal hub block assembly which provides transmission of drive torque from the transmission assembly output to the rotor hinge assembly and rotor blades via attachment lugs; PA1 (2) a mechanical hinge assembly of machined metal components, pins, and bearings which secure the blades to the rotor hub while permitting flap motion, lead-lag motion and feathering motion; PA1 (3) a flexible mechanical assembly of metal straps or wire strands which transfer centrifugal force load from the blade root end to the hub block with minimum feathering resistance; and PA1 (4) a lead-lag damper.
A current metal hub rotor system manufactured by the assignee of the present invention for the CH-47 helicopter is a typical example of the traditional articulated rotor systems noted above. Historically, few problems have been encountered in achieving satisfactory static strength in such a rotor system. The problems that were encountered centered mainly about component endurance in terms of fatigue life and wear. The fatigue life has been improved somewhat by improved design materials and processes. The wear problem is manifested mainly in bearing rolling elements, races, and seals. Bearing design has improved through practical experience and analytical and test developments. Seal integrity has been improved, and grease lubrication has been replaced by oil lubrication. Further, metal alloying, heat treatment, machined surface finish and surface hardening and cold working methods have also been improved.
There is a limit, however, to the traditional rotor system technology improvement cycle using known materials and techniques. This limit is of particular concern when it is noted that the growing necessity to increase aircraft speed and to reduce fuel consumption requires reduction in hub drag which limits bearing and hub component size, but does result in a weight reduction. The weight reduction means less load-carrying material, hence higher stresses.
Rotor system design technology, in this dilemma, turned to utilization of new metal alloys, reduction or elimination of bearings and introduction of new design concepts. Titanium alloys have been used to replace steel in more recent designs, benefitting somewhat from improved specific fatigue strengths. Specific ultimate strength and specific stiffness, however, are similar among the leading metal candidates, i.e., aluminum, titanium and steel. When cracks are initiated in the notch sensitive titanium alloys, crack propagation rates are of vital concern.
The designers next turned to non-metallic materials such as elastomeric materials for bearings and fiber reinforced materials (composites) for structural members, The advent of these non-metallic materials has opened a new avenue of innovation and improvement in aircraft design.
The application of composite materials to rotor systems indicates, for example, that they have potential in providing a solution to many of the ailments of metallic rotor hubs. Composite materials have been shown to improve life, damage tolerance and fail safety due to their relative notch insensitivity, slow crack growth, superior fatigue strain endurance and high strain energy storage prior to fiber failure. The raw material is basic in that it can be sized and shaped to any proportions with a minimum of trim and scrappage, allowing strength and stiffness to be discretely introduced only where required. Techniques have been developed to reduce damage propagation even further through material hybridization and fiber and layer orientation. Composites are not susceptible to corrosion and are readily inspectable using ultrasonic and/or radiographic techniques to detect flaws in their laminar and fiber content. Their high specific strength and stiffness offer the potential of significant weight reductions.
The application of composite material to a main rotor hub and hinge assembly of a helicopter has been suggested. See, for example, U.S. Pat. Nos. 3,762,834 and 4,012,169. These so-called "star-like" structures are essentially formed as continuous plates with three extending arms. In U.S. Pat. No. 4,012,169, each arm comprises three parallel plates for providing a truss configuration. These, however, have been applied to single main rotor helicopters and not to tandem helicopters. The two systems (single rotor and tandem) do have differences. For example, a tandem system may require a higher lag hinge offset than that resulting in single rotor systems because lag deflection of the blade tip, for the condition of CF and drag movement equilibrium about the lag hinge, is inversely dependent on the radial placement of the lag hinge. It can be concluded, therefore, that if the design criteria peculiar to a tandem system are met, then the resultant configuration can be applied directly to a single rotor helicopter. However, if only those goals peculiar to a single rotor system are achieved, then the resultant hub, for example, would be unsuitable for application to a tandem system.